Mars Direct; Mars Semidirect; Design Reference Mission. Need Updating?

SpaceNut · 2017-02-25 23:01:00

While doing some more searching the web I did come across this https://solarsystem.nasa.gov/docs/Bayle … -Paper.pdf
Page 3 has a graph of the atmospheric density to altitude and it seems that its pretty solid and consistant at 30km and down but above this its seems to vary quite a bit.

RobertDyck · 2017-02-26 02:33:06

GW Johnson wrote:

Somewhere, somebody has to address the wisdom of exploring one site versus many. 500 years ago it was many, for the most successful ventures of that time. There's real lesson about what "exploration" really is there.
If you go with many sites, all this minimalist throwaway stuff is nonsense. You have to start thinking recoverable and reusable designs, for as much of the architecture as is humanly possible. Use it many, many times for multiple missions and purposes, and in spite of the higher thrown weights, it becomes ultimately cheaper, but only as amortized over a life cycle.
GW

Zubrin had argued for exploring multiple sites before settling on one to build a permanent base. He argued exploration sites should be within rover range of each other, so Mars explorers could go from one former Mars site to another. He called them "warming huts". I have argued that was fine in 1990, but NASA has spent 27 years and counting doing exploration with robotic vehicles: orbiters, landers, rovers. All that time and money has to be good for something. Instead settle on one site now, start construction of the permanent base with the first human mission. That doesn't mean you're restricted to one site. Initially use the rover to explore many km from base. Eventually use a "hopper" rocket to travel vast distances from base.

I also argued to make the Interplanetary Transit Vehicle reusable. To travel from ISS to high Mars orbit and back. Aerocapture at each planet. And the surface habitat will be reusable, each mission building up the permanent base. Initially with expendable propulsion stages, but when technology becomes available, those stages can be replaced by a reusable one.

SpaceNut · 2017-02-26 09:56:51

Entry, Descent & Landing Sequence Overview of MSL
This has lots of details for the timing and such for GW to read through for retro propulsion useage.

Working backwards from the numbers for the permanent base once we establish those numbers for what is that mass we then work backwards towards what mission archecture can deliver what we need and when.

If all you want is empty room area send down an Inflateable and hope that you can move all you need into it later. Otherwise send down what many call tuna cans and be done with the construction other than expanding what we have.

SpaceNut · 2017-02-26 10:17:28

Origin of the Apollo-shaped Manned Mars Lander (1966)

Woodcock’s 56.1-ton MEM would comprise a descent stage roughly 33 feet across (the diameter of a two-stage Saturn V rocket) at its widest point and, hidden beneath a protective nose-cone (“separable cap”), a 27.3-ton ascent stage “payload.”

Woodcock envisioned that his MEM would form the basis of a long-term, increasingly capable and complex Mars exploration program. He proposed a design for a one-way logistics lander in which cargo and a “camper-type” pressurized rover would replace the MEM ascent stage and the surface operations shelter.

Crew support:

Woodcock calculated that 10.6 tons of water, food, and oxygen with a four-ton reserve could sustain a five-man crew in the MEM on Mars for 500 days. Like the logistics MEM, the power and shelter MEMs would land on Mars unmanned.

Oldfart1939 · 2017-02-26 13:35:39

I've been doing calculations since yesterday evening using the so-called "Rocket equation," and have come up with what I call my initial Mars mission. This involves a much enlarged and modified Dragon 2 spacecraft, one with an integral cargo trunk section modified for pressurization and habitability, including a meal area, sleeping accommodations, and exercise equipment. Plus food storage. I've modified my architecture to a crew of 5. The dry mass of this portion of the system is ~ 35,000 kg. This is all within the ability of the Falcon heavy to make an orbital throw. This would be coupled with a 114,000 kg modified Falcon 2nd stage--modified to fuel by methylox propulsion and the new Raptor engine. Dry stage mass of 6500 kg, and 107,500 kg combined LOX and LCH4. These second stages would be equipped with extensible landing legs, similar to Falcon first stages currently in use. This system should be capable of meeting or exceeding the delta V of 4000 m/s required for trans Mars trajectory. There would be sufficient fuel remaining in order to land propulsively with considerable excess on hand. This vehicle could be refueled by ISPP and return to Earth orbit using aerobraking. This also calculates as "possible" using the Merlin engine and RP-1, but whether the RP-1 would remain fluid is highly doubtful? My other option is launching these systems 2 at a time--in parallel in order to use a tether between spacecraft and mutual generation of artificial gravity. The only real hangup now is getting the Falcon 2nd stages to LEO and subsequent docking with the spacecraft. I also propose they be capable of docking at the ISS through the standard cargo hatch; the spacecraft could be flown to orbit unmanned, and crew temporarily stationed at the ISS. It's conceivable that these second stages could be recycled in this manner and there are few if any throwaways. If sufficient fuel remains on board after LEO return, the possibility of a hypersonic retropropulsive atmosphere entry recovers these stages?

I've been mulling this over for a long time. Naturally, there would be several missions on order to accumulate the necessary base, preposition landing transponders, and transport the ISPP "factory" along with the Nuclear reactor and habitat construction materials.

SpaceNut · 2017-02-26 16:20:57

Some of the rocket equations that I have seen are just something that I can not seem to grasp. Is there one you might recommend as calculus is out of my scope of capability at this time.

Oldfart1939 · 2017-02-26 16:31:56

Robert Zubrin does an excellent job in his second book: "Entering Space; Creating a Spacefaring Civilization." A brief discussion of principles of rocketry on pp. 35-38 shed a lot of light on things. I had a lot of math--55 years ago--mostly forgotten by now. I spent several hours last night working on several derivations of Differential equations that I couldn't solve anymore. So...thanks to Bob Zubrin, it is now making some sense. I'm also hoping for GW to take a look at my conclusions. I hope I'm correct, or not too wrong!

GW Johnson · 2017-02-26 17:21:31

Oldfart1939:

My best calculations show a worst-case 3.94 km/s to depart for Mars on a min-energy Hohmann trajectory, duration about 240 days. If you do nothing, and if you do not miss the mark at Mars in some way, then free entry at Mars is just above Mars escape at 5.04 km/s. This needs to be a very shallow entry angle so as not to strike the surface while still hypersonic. That situation is entirely unlike warhead entry here on Earth. Unfactored for anything.

Otherwise, the delta-vee to enter low orbit about Mars is right at 1.59 km/s, unfactored for anything. The value of the mid-course corrections required not to miss the mark at Mars in either case is something I cannot estimate with pencil and paper. I would hazard the wild guess it is 0.1 to 0.2 km/s.

On the other hand, if launched direct from Earth, your rocket needs to be capable of at least 12.14 km/s, assuming gravity and drag losses are 5% of the circular orbit portion of the delta-vee. That value is factored by 1.05 on the theoretical circular orbit portion, because surface escape (the "theoretical" value) is 11.18 km/s. The theoretical surface circular orbit value is escape divided by square root of 2. Use it to represent circular orbit at low orbit altitude, where orbit speed is really slightly slower than the surface value, the difference being attributable to potential energy at that altitude.

These velocities combine not as velocities but as kinetic energies. You need a velocity "at infinity" from Earth that is the difference Vinf = Vper - VEarth between Earth's orbital velocity and the perhelion velocity of the transfer orbit. It combines with Earth's surface escape velocity (which builds-in the potential energy) as dV =(Vesc^2 + Vinf^2)^0.5. That theoretical gets split into dV - Vcirc + Vcirc, and you apply your gravity/drag factor to Vcirc, the surface circular orbit velocity. Then add some more for undrainable propellant, and some more still, for midcourse corrections.

The rocket equation itself is quite simple: dV ideally delivered = exhaust V * LN(MR) where MR = Wign/Wbo, the mass ratio for the stage burn. You can estimate exhaust velocity pretty closely as (Isp, sec) times g (either 9.8067 m/s^2 or 32.174 ft/s^2), but only if your Isp is appropriate to the actual burn conditions. That's where most people make their mistakes; estimating the right Isp really requires doing realistic engine ballistics, starting with characteristic velocity c* as a function of chamber pressure and mixture ratio. Most of the reported Isp data people are using are too high, being this or that definition that is simply not so realistic.

This ideally-delivered dV must cover the theoretical orbital mechanics-derived dV plus any gravity and drag losses (around 5% to orbit at Earth), plus any undrainable propellant fraction from the tanks (typically near 2%), plus any course correction or maneuvering delta-vee budget (fairly low but who knows????). So factor up theoretical dV enough to cover those things before you solve the rocket equation for the MR.

Use a total of maybe 10% beyond orbital mechanics theoretical values, and you should be in the ballpark. For a given value of MR, the propellant mass fraction is 1 - 1/MR. Payload fraction plus propellant fraction plus total-of-inert structures fraction must be 1. If there's not enough payload for a given level of inerts, payload will show as negative. That says the design is totally infeasible.

GW

Last edited by GW Johnson (2017-02-26 17:53:40)

Oldfart1939 · 2017-02-26 17:43:32

GW-

My calculations were for a delta V of 4.0 km/sec departing LEO; there's some wiggle room should a shorter time trajectory be desired. I also included a massive surplus of food and water in my Spacecraft mass. There's a full 30 month food supply calculated in case of an Earth Free return trajectory. It is doable with an RP-1 fueled rocket, but I'm worried about the RP-1 turning into a hydrocarbon equivalent of Jello. The step up to Methane helps a LOT w/r to Isp available and a much higher exhaust velocity in the Rocket Equation calculations. I should probably recalculate on the basis of a delta V of 5.0 km/sec departure, which should give a E-M transit time of 180 days, but still with a free return available. I'm just wondering what Elon Musk's next surprise will be? It would make sense to me for the heavy satellites he's contracted to place in GST for an upgraded Falcon 9 second stage in order to demonstrate the Raptor engine "in flight," and also showcase the Methyoxy fuel oxidizer couple. Your thoughts as to how much that could improve the lift of mass to LEO?

You added something in an edit while I was doing my last post. Here are the numbers I used: Base mass of the spacecraft minus fuel; includes the 35,000 kg and 6500 kg for the Falcon 9 rocket second stage dry weight. Then: 107,500 kg combined methane and LOX. Isp for CH4 (vacuum ):348 sec. A deltaV/C = 1.173; e^1.173 = 3.23 for the mass ratio (min). Mass ratio (actual) = 3.59. So...I have some wiggle room for a possible Mars landing, if not with the whole works but the crew capsule alone.

Last edited by Oldfart1939 (2017-02-26 18:01:07)

GW Johnson · 2017-02-26 18:03:56

Well, there's improvement, but it's not all that dramatic. The 1000 psia chamber c* data in my old Pratt & Whitney handbook are very little different between CH4-LOX and kerosene-LOX. The difference between kero-LOX and NTO-hydrazine is not so very much either.

The real difference between kerolox and methane-lox is in the reusability. It is the fuel that cools the engine inside the regenerative passages. Kerosene tends to coke in these passages when it gets hot; methane, not so much. A big problem refurbishing for the next flight is cleaning out all the coked carbon in the passages.

The problem with cryogenics at all is boiloff over long periods of time. The problem with kerosene is it gels cold, unless you keep the tank above -50 F. I think you also have to keep hydrazine and NTO from freezing. So there's just not that much difference.

The only other difference is ignition. You need a vacuum-rated igniter for kerolox or methane-lox. NTO-hydrazine is hypergolic, even in vacuum.

GW

SpaceNut · 2017-02-26 18:35:48

I typically will use comparison scaling when I look to what we have and what we want to do with it.

So lets look at the Falcon 9 heavy which can lift 53 mT to earth orbit and when I look at the same rocket being launched from mars surface we can scale it down by a factor of 2.66 as mars is only 0.375 that of Earth for gravity.

Or even think of it this way when we launch a falcon 9 v1.2 FT we can loft to orbit 22.8 mT so if this was setting on mars and scaled down by the 2.66 factor then we have a payload of ~8.6 mT which is approximately the size of a red Dragon. Normally its 70 m tall for payload and both stages. Which suggests a scaled down height of ~26.5 meters which is a long way to the ground from a capsule door but that could change as we work through the stage numbers. The Dragon Spacecraft & Trunk 8.1m (26.6 ft) height. Payload flairing is 5.2 m tall but I think we can use the dragons height once we recalculate the stages under the capsule changes.

Since Mars is a near vacuum it only makes sense to configure the lander with the correct engine (934 kN (Vacuum) Merlin Burn time of 397sec) for this iteration of design meaning we need for the second stage only ~351 kN for mars. Which means only 3 engines are required.

The normal first stage engine at sea level, for a total thrust of 6,804 kN ( Merlin 1D burn time 162sec) means we need to target ~ 2,560 kN for launch from mars surface for 1st stage. One engine is only 756 kN so giving the new first stage the 934 kN (vacuum engine) only makes sense.

Current Space x page indicates first stage Thrust At Sea Level 7,607 kN Thrust In Vacuum 8,227 kN, which appears to be 47 m tall. Mars height would be 17.7 m tall for the stage. Normally both stages plus flairing is 70 so if we remove the Pay load flairing height of 5.2 m out of the 70 m and 47 m first stage then the second stage is 17.8 m tall for the falcon 9 rocket and for mars at ~ 6.7m.

Merlin Vacuum Certification On March 7, 2009, SpaceX performed a full mission duration firing of the new Merlin Vacuum engine at McGregor. The engine fired for six minutes or 360 seconds, consumed 45.36 tonnes of propellant, and demonstrated a vacuum specific impulse of 342 seconds, highest ever for a U.S. hydrocarbon rocket engine.

The two-stage, 313 tonne, kerosene/LOX rocket so proportional for mars means ~118 t of fuel and oxidizer.
The second stage normally fires for 397 sec. so probably close to 50 t for the stage and for mars its 18.8 t.
The first stage then calculates to 263 t of which for reusuability is about 70 % leaving the rest for landing. So if we were on orbit and coming down then we would need ~79 t so for a retropropulsion we need ~ 30 t to land not count stage and capsule mass plus on orbit.

So mars sized falcon 9 is 32.5 m tall for a dragon plus truck of 8.1 m with a first stage 17.7 m with fuel/oxidizer of 30 t with a second stage that is 6.7 m tall and fueled with 18.8 t of which we would need to land with the second stage empty and a bit large fuel fill for the first stage for th combined landing attempt.

Seems like we could make this a single stage to orbit with the 9 vacuum engines.

Of course as meantioned the RP-1 and Lox is not the correct fuel for mars if we are making insitu fuels so now we would redo these calculations for the new fuel type and engines that we could use.

Something else is to change the 3.7 m diameter to the more practical 10 m to make the stages shorter as well.

With this long post we can see that for mars we are only landing a stackup of no more than the hieght of the current stage Falcon 9 for the entire mars lander. Which is a small 40 m tall at the 3.7 m diameter but if we go to the 10 m diameter then we are looking at something under 15 m tall. That said 4 engines for a complete single stage to orbit would seem to be correct.

SpaceNut · 2017-02-26 20:56:31

So now to covert for the methane-lox and engines for the same scaling once engines numbers are looked up.

While we could land with NTO-hydrazine we do not have an insitu ability developed yet for manufacturing it for refueling.

The same holds true of NTO / MMH even with the super Draco engines.

We really need the ability to make other insitu fuels so as to be able to use what we have rather than changing plans for what we need developement for.

https://en.wikipedia.org/wiki/Dinitrogen_tetroxide

Nitrogen tetroxide is used as an oxidizer in one of the more important rocket propellants because it can be stored as a liquid at room temperature. Nitrogen tetroxide is made by the catalytic oxidation of ammonia: steam is used as a diluent to reduce the combustion temperature. In the first step, the ammonia is oxidized into nitric oxide:
4NH3 + 5O2 → 4NO + 6H2O
Most of the water is condensed out, and the gases are further cooled; the nitric oxide that was produced is oxidized to nitrogen dioxide, which is then dimerized into nitrogen tetroxide:
2NO + O2 → 2NO2
2NO2 ⇌ N2O4
and the remainder of the water is removed as nitric acid. The gas is essentially pure nitrogen dioxide, which is condensed into dinitrogen tetroxide in a brine-cooled liquefier.

https://en.wikipedia.org/wiki/Monomethylhydrazine
Monomethylhydrazine (MMH) is a volatile hydrazine chemical with the chemical formula CH3(NH)NH2.

Oldfart1939 · 2017-02-26 21:43:16

I also looked into using the Russian Proton M third stage, since it could be boosted into LEO by Falcon Heavy in the presently as designed configuration. It is NTO/ADMH fueled but there isn't enough fuel available for a TMI. The Isp is fine, but the amount of fuel comes up way short. I didn't do the calculations for staging 2 of them in sequence, however.

Last edited by Oldfart1939 (2017-02-26 21:48:12)

SpaceNut · 2017-02-26 23:22:36

The trouble with looking to get a comparible sizing for Space x is that they have not sent any thing as of yet to Mars only the Atlas V has done so. https://en.wikipedia.org/wiki/Atlas_V
The Atlas V -541 has been used and will be used again for mars. capable of
It uses the American-built RL10 engine burning liquid hydrogen and liquid oxygen to power its Centaur upper stage.
Powered by either one or two Aerojet Rocketdyne RL10A-4-2 engines, each engine developing a thrust of 99.2 kN.

Second stage - Centaur
Length    12.68 m (41.6 ft)
Diameter       3.05 m (10.0 ft)
Empty mass 2,316 kg (5,106 lb)
Propellant mass 20,830 kg (45,920 lb)
Engines    1 RL10A or 1 RL10C
Thrust    99.2 kN (22,300 lbf) (RL10A)
Specific impulse 450.5 seconds (4.418 km/s) (RL10A-4-2)
Burn time 842 seconds (RL10A-4-2)
Fuel    LH2/LOX

GW Johnson · 2017-02-28 12:36:09

Hmmm, very interesting that the 1966-vintage concept for a 2-stage Mars lander looks so very much like the single-stage and 2-stage designs I have been looking at. Spacenut put those in post #29 above. I had actually never seen those sketches. Thank you, Spacenut. I have no idea where you dig up stuff like that.

With 2-stage you can carry a lot of tons down, even if you do a full retropropulsion landing that kbd512 hates so much. The descent delta-vee is only on the order of 1 km/s. Only a little bit of tonnage (crew plus samples) makes the return to orbit, theoretically 3.6 km/s, probably reliably closer to 3.7 or 3.8 designed-in.

Single stage probably needs about 4.8-4.9 km/s capability designed in. That pushes the MR very much harder on descent, so that cargo tonnage is either quite small, or your vehicle is quite huge, or some of both. Depends almost entirely on what inert fraction you think you can get away with, not so much on propellant choice. Heat shield isn't the bugaboo so very much, it's landing legs and loading ramps and aeroheat backshell that get heavy really quickly, especially if you intend to fly more than once (which requires that you retain them). But it really is doable.

GW

Last edited by GW Johnson (2017-02-28 12:39:23)

SpaceNut · 2017-02-28 21:25:45

I actually was part of another group that was working on a common core lander to which the shapes of the sketches do show. They also show why the base diameter needs to be closure to that 10 m and not the 3.7 m of the space x rockets.
I think that I am pretty close to a mini falcon rocket design for mars as a lander in post #36 but it needs to be tweaked for the fuel that we can make on mars or look for a total shipment from earth for refueling it to make it reuseable which is not realistic.
Scaled testing of raptor and not a full size means that we need to wait for those to be integrated into any concept of use.

kbd512 · 2017-02-28 22:37:43

GW,

If I told you a new $10 wrench would cost $10K to deliver and that you must have a new wrench delivered after every instance when the wrench is used, how many wrenches would purchase before you started questioning the utility of the machine that required the wrench? Pure propulsive landings dictate that if I want to deliver something, then I must first devise a way to deliver tens of tons worth of propellants and aerospace vehicle.

I'm not opposed to retro-propulsion if it's necessary, as is the case for places like our moon. I'm just opposed to the insane cost associated with sending many tons worth of rocket and propellants to another planet with an atmosphere. There's very little contributed to the mission by delivering dead mass. After a rocket lands, it's a paperweight if you can't refuel it. There are no propellant plants, launch pads, or other infrastructure of any kind on Mars. That's the crux of the problem. Apart from whatever is required to return humans to Earth, for exploration purposes, anything we deliver to Mars should stay on Mars and provide utility to the people we send it with while they're on Mars.

I want humans to go to Mars to produce scientifically valuable information about Mars and to search for signs of past life. We've done what we can reasonably do with robots and we still have a lot of unanswered questions. Locating water and useful minerals or ores to assist future colonization efforts are secondary objectives.

There is no reality-based requirement for reusable multi-stage landers for us to achieve those objectives. I can all but guarantee that pure propulsive landers will simply make this exploration target so expensive to explore that we simply will not go. New thinking is required here.

How do we soft land on Mars with minimal complication and expense? We already know how complicated and expensive pump-fed liquid fuel rockets are, so unless it's absolutely required that can't be the answer.

GW Johnson · 2017-03-02 14:51:49

Kbd512:

It's just not that stark a choice between all-aerobraking and aerobraking-with-retropropulsion for landing. From orbit, retropropulsion is roughly 1 km/s out of 3.6 versus ~0.2 km/s out of the same 3.6 for chutes with a last-second retrorocket touchdown. Yeah, the chutes help, if you can make them work. I base that on less-than-half-a-Mach terminal speeds in that thin air.

To make chutes work with large objects will require the inflatable or extendible heat shield ideas to get the out-of-hypersonics altitude way up. These technologies show great promise in tests so far, but we have as yet no experience in actual practice. Bureaucracies as unwieldy and risk-averse as NASA will use that inexperience as another excuse not to go.

The relative difference in the retropropulsion deceleration numbers is even smaller if you instead consider direct landing: 1 km/s out of maybe 6 km/s for full retropropulsion / no chutes, versus 0.2 km/s out of that same 6. It's just not that big in comparison to the aerobraking that you can do, either way.

It really doesn't look like the killer you indicate, to me, especially for one-way stuff. It's that 2-way "Mars ascent vehicle" stuff that's the killer. You just have to grit your teeth and do it. Except it's just not that bad.

To reach Mars orbit from the surface requires a designed-in delta-vee at or exceeding 3.7 km/s. Even in a two-stage design, that's tough enough for design, but it's quite far from the edges of technology. With vac-design nozzles and NTO-MMH storables (the lightest tank designs), MR = ~3.08, which is 68% propellants. That assumes I achieve 3.28 km/s exhaust velocity from 335 sec Isp.

How many tons returns is basically what sets the mass of the ascent vehicle. For say 5% inert weights, and only 3 tons sent back up to orbit, the ascent vehicle is near 11 or 12 tons at ignition. It's then in turn the payload for the descent stage, which needs perhaps 10% inerts to account for heat shield and landing legs, and the structural strength to hold the ascent stage through all the flight and landing loads. This sizeout is ~8 tons ascent propellant.

Now on descent, MR = 1.36 for 26% propellants at dV = 1 km/s, and MR = 1.06 for 6% propellants at 0.2 km/s. That difference of 10% gets reduced some by the weights of extendible or inflatable heat shields and the weight of the chutes. For a no chutes rocket landing, carrying a 12 ton ascent vehicle, your package masses about 33 tons at entry, with about 8 tons of descent propellants. For a chutes/extendible heat shield landing, your package masses about 14 tons at entry plus the mass of all the chutes and extendible heat shield gear, and has about 1 ton of descent propellant. A decent guess: 16 tons at entry? Can all this extra aerobraking gear fit in a 2 ton budget?

The differences are only 17 tons at entry, and only 7 tons of propellant. That and the fact that retropropulsion is already being used by Spacex and Blue Origin, and also decades ago by Grumman on the moon. We already know it works just fine. The extendible heat shields are not yet being used. Hopefully they soon will be, but I see no signs of it yet.

16 vs 33 tons at entry to raise 3-4 men in suits and a hundred kg of samples back to orbit is significant, yes. A program killer? I doubt it. That's an entry mass factor around 2, not near 10! I do not deny that every ton sent to Mars requires several tons sent up from Earth. But at only factor 2-ish, that's only a problem if you lock yourself out of orbital assembly using inexpensive commercial launchers, which SLS will never be (nor will you ever have many around to use in a single year).

Same picture still holds when you rescale to other sizes (assuming anybody disagrees with my 3 tons of people in suits and samples plus minimal life support for maybe 2 days), and it doesn't shift all that much if you try to design-in some on-orbit maneuvering into the ascent stage. That's just a little bit bigger MR and propellant percentage to deal with.

It gets even better with methane-LOX versus the NTO-MMH, which is what I assumed. If you can actually make rocket-grade material quality in the dozens-of-tons quantities we need, and in the timescales we need (something I think must be demonstrated BEFORE we go), then you can land with your 8 tons of ascent propellants deleted in favor of 8 tons of descent cargo! For the silly example given above, that's about 8 extra tons of stuff delivered to the surface of Mars. Either way, this is because it's the ascent propellant we make, not the descent propellant, which is much smaller (8 versus 1 ton).

The tankage won't be quite as light with methane-LOX cryogens, because you'll need more insulation than the storables need. This affects the inert fractions a tad. A bit of insulation and a small tank heater is all the storables really need.

GW

Last edited by GW Johnson (2017-03-02 15:51:15)

Oldfart1939 · 2017-03-02 15:29:55

GW-
Looking at your numbers really points out why using ISPP is nearly essential to success of a Mars landing and Earth return. The mission architecture I espoused elsewhere concentrated on using an entirely methylox set of vehicles for that reason alone. Yes, MMH and NTO are very advantageous for really long term storage. What is the boil off rate of LOX from a deep space vehicle, anyway? LCH4 should be less problematic.

GW Johnson · 2017-03-02 15:53:58

Hi Oldfart1939:

To answer your question about boiloff rates of cryogens, I honestly don't know. I just know it gets bad with simple uninsulated tanks after a very few hours. Insulated tanks with sun shields and cryocooler rigs will be heavier. No way around that.

I rather think that you can use LOX-LH2 to depart Earth, but you'd better be using NTO-MMH to arrive and to land, because of the months in space. You can use LOX-LCH4 to ascend, but your return vehicle had better be using storables if parked in orbit, because of the months to get there, and the months waiting while the crew is on Mars. Otherwise, you must ship many, many tons of LOX-LCH4 up from Mars to power it. If you recover in LEO, the months of travel make storables the logical choice there, too.

Musk sidesteps this by landing his giant ship on Mars, and refueling it there. The downside is that the single stage craft must escape from Mars and recover at Earth. That mass ratio drives his design into carbon-epoxy structures, even for the LOX tankage. Dangerous!

He gets around the confined space problem by that same choice to land his spaceship directly. You can ride a cramped capsule up to Mars orbit, but you had better have a spacious habitat for months it takes to get home. We already know from Gemini 7 that ~3 weeks in a cramped capsule edges very close to insanity. It takes 6-9 months to come home from Mars.

GW

Last edited by GW Johnson (2017-03-02 16:06:26)

Oldfart1939 · 2017-03-02 16:06:53

Well, a lot depends on the ambient temperature and thus, rate of heat transfer to the cryogen tank. In deep space, I wouldn't expect the rate to be very high, provided the tank be shaded from the Sun. As always, "the devil is in the details."

Oldfart1939 · 2017-03-02 16:24:30

I was initially trying to use existing hardware for everything in my modelling, but as usual, mass is the killer. I was looking at the availability of MMH/NTO fueled vehicles, and there are some components in the Russian Proton M that might have some promise for incorporation. For an EDS stage, the SpaceX Falcon 2nd stage could be used, but would be difficult to get to orbit; if it were only partially fueled, it would still be able to provide a big delta V to my 35,000kg ship, which now would need to incorporate a MMH/NTO fuel system and motor(s). The ERV would need to be ISPP to LMO, and meet a fully fueled Proton M 3rd stage and 4th stage for TED (Trans Earth Departure). Alternatively, a truncated falcon second stage with reduced fuel capacity and lower mass could be used for the EDS. Back to the drawing board (and calculator!).

SpaceNut · 2017-03-04 21:31:48

The fuel type seems to be problematic to making what we have to making it work in an altered state.

If we have some numbers for the space x second stage gelling of fuel due to wait times in orbit then we would know how to heat the fuel to keep it at the right consistancy. With that we could stack several of them on a seperate Falcon heavy to be sent for sub-assembly at the ISS and then we have the EDS. They already have the matching interchange coupling collar to make use of whne putting them together.

kbd512 · 2017-03-05 01:24:15

GW,

IRVE-3 used rotation for stability and cold-gas RCS for attitude control at the entry interface. Spinning humans may not work so well. MSL used spin for stability and RCS for attitude control. Could spin negate or lessen the RCS requirement, or is that a bad idea?

A new method for HIAD control NASA is working on is called the Cable-Controlled Aeroshell Deceleration System (CCADS). It is intended to provide coarse and fine HIAD flight control during descent by varying L/D to adjust cross-range, a concept much like using the body flap on the orbiter to adjust aerodynamic lift (CCADS basically uses cables to "pinch" the HIAD):

CCADS Presentation

CCADS Technical Paper Final Draft

The "HEART Flight Test Overview" report said shallow angles of attack (10 degrees or less) negated the requirement for thermal protection of the exposed portion of the Cygnus. Is that an error with the Thermal Desktop heating model used to determine peak heating or is that possible? The CCADS flight profile uses entry angles greater than 10 degrees. Is it possible to use a 10 degree or less followed by actuation of CCADS to increase lift and establish a more horizontal glide path to the ground?

The computer models indicate that HIAD (no parachutes or CCADS) is subsonic at 15km, but then velocity is nearly constant between .3km/s and .2km/s all the way to the ground. There is virtually no deceleration from .2km/s at approximately 5km. Apparently, high subsonic flight speeds are required to generate lift. CCADS lengthens flight duration considerably, permitting astronauts to actually "fly" Cygnus to a landing target, but it's either traveling at .2km/s or it ceases to generate lift and subsequently falls out of the sky.

A subsonic parachute and mortar system would weigh more than rocket engines and propellant with .25km/s dV. I give up. There's no better way to do this.

My proposed propulsion module is mounted atop Cygnus over the docking ring since HIAD is on the bottom. It has 8 MR-80C's (AeroJet-Rocketdyne MR-80B's modified to use HAN / AF-M315E monopropellant) for retro-propulsion and 16 MR-104D'S (AeroJet-Rocketdyne MR-104C's modified to use HAN / AF-M315E monopropellant) for attitude control. Isp (vac) for MR-80C and MR-104D is presumed to be 250s. AF-M315E's density is 1.47g/cm^3. A US gallon of AF-M315E weighs approximately 12.26774lbs, so 450kg of the stuff is slightly less than 81 gallons. My retro-propulsion engines are radially mounted over the docking ring in a sort of "nosecone" for the Cygnus PCM (a misnomer since Cygnus is traveling base forward throughout EDL).

Notes:
1. Cygnus won't fly aboard Antares, thus no requirement for the structure to withstand 8.5g peak acceleration, so it'll be a bit lighter
2. HIAD is separated just prior to retro-propulsion and Cygnus becomes 1,000kg lighter
3. RCS module is permanently attached to Cygnus, but hinged so the crew can ingress / egress through the PCM's top hatch
4. Cygnus is modified with the addition of a bottom hatch for normal ingress / egress on the surface of Mars
5. Solar panels for surface power are stowed inside Cygnus and manually attached by the crew to the top hatch
6. Cygnus landing gear consists of composite struts with small wheels containing electric hub motors for 2.5km/h max speed
7. Cygnus lands several kilometers from the ascent stage to avoid potentially damaging the ascent stage during retro-propulsion
8. Consumables mass assumes the crew consists of 1 average man and 1 average woman
9. Service Module provides power and propulsion for transit
10. Apart from providing transit power and mid-course correction burns, the primary purpose of the service module is to decelerate the PCM approximately 1km/s just prior to reentry, at which point in time it detaches from the PCM
11. My previous mass estimates for a scaled-up HIAD were used for the ascent stage. Mass was greater because the ascent stage has a greater mass. I mistakenly added that mass to the Cygnus reentry mass. My Cygnus PCM is within the 5600kg limit imposed by the HEART HIAD precursor mission design.
12. Cygnus is positioned atop the ascent stage
13. Ascent Stage uses 4 AeroJet-Rocketdyne AJ10-190's (the uprated OMS-E / OME engine developed for Orion); essentially a reprise of MIT's Scott Alan Geels Mars Ascent Vehicle, 4.2km/s dV (accounts for drag and gravity losses); requires a 10% or less inert mass fraction; nearly maxes out Falcon Heavy's throw capability, requiring 13t to TMI
14. Earth Return Stage uses 1 AeroJet-Rocketdyne AJ10-190 (the uprated OMS-E / OME engine developed for Orion)
15. Consumables and HIAD #2 for Earth return are delivered with the Earth Return Stage (some assembly required)

Mass Estimates:

Service Module (1 AeroJet RocketDyne AJ10-190): 4,750kg
Gross Reentry Mass: 5,250kg
HIAD: 1,000kg
Cygnus PCM gross mass (reentry; includes RCS module): 4,250kg
Cygnus PCM (structure, avionics, life support - CAMRAS / MOXIE / IWP, batteries, solar arrays for surface power): 1,650kg
Consumables (food and water): 1,850kg
Astronauts (we're sending small people to Mars): 150kg
RCS Module: (450kg HAN, 68kg for 8 MR-80C, 30kg for 16 MR-104D, 52kg for structures): 600kg

Well, there it is. It's quite minimalist, but it should get the job done. I would feel better about this if there was another Cygnus loaded with consumables and parked in the landing area. A Mars orbital station like Lockheed-Martin's Mars Base Camp would be nice, too. All the Falcon Heavy payloads range from 11t to 13t. If I had 15t to work with, the mission hardware elements could be more robust. Unfortunately, 13.6t is all Falcon Heavy can deliver. C'est la vie.

SpaceNut · 2017-03-05 21:47:10

The AeroJet-Rocketdyne 8 MR-80B and 16 MR-104D happens to be the same as the landing for the 2020 rover which was the same as the curiosity rover as well using the skycrane landing cable system.
http://www.rocket.com/article/aerojet-r … er-mission

The cygnus still used a payload shroud even when launched on the Atlas V so it will still need one for use to get out of earth atmopshere but it will still jetison it for a mars landing since the cygnus will use a heatshield and HIAD for getting into mars landing profile.

Quite a fact filled post so still reading and digesting its content....

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